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Today, the near‐Earth space is facing a paradigm change as the number of new spacecraft is literally skyrocketing. Increasing numbers of small satellites threaten the sustainable use of space, as without removal, space debris will eventually make certain critical orbits unusable. A central factor affecting small spacecraft health and leading to debris is the radiation environment, which is unpredictable due to an incomplete understanding of the near‐Earth radiation environment itself and its variability driven by the solar wind and outer magnetosphere. This paper presents the FORESAIL‐1 nanosatellite mission, having two scientific and one technological objectives. The first scientific objective is to measure the energy and flux of energetic particle loss to the atmosphere with a representative energy and pitch angle resolution over a wide range of magnetic local times. To pave the way to novel model‐in situ data comparisons, we also show preliminary results on precipitating electron fluxes obtained with the new global hybrid‐Vlasov simulation Vlasiator. The second scientific objective of the FORESAIL‐1 mission is to measure energetic neutral atoms of solar origin. The solar energetic neutral atom flux has the potential to contribute importantly to the knowledge of solar eruption energy budget estimations. The technological objective is to demonstrate a satellite deorbiting technology, and for the first time, make an orbit maneuver with a propellantless nanosatellite. FORESAIL‐1 will demonstrate the potential for nanosatellites to make important scientific contributions as well as promote the sustainable utilization of space by using a cost‐efficient deorbiting technology.
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manuscript submitted to JGR: Space Physics
FORESAIL-1 cubesat mission to measure radiation belt
losses and demonstrate de-orbiting
M. Palmroth1,2, J. Praks3, R. Vainio4, P. Janhunen2, E. K. J. Kilpua1, N. Yu.
Ganushkina2,5, A. Afanasiev4, M. Ala-Lahti1, A. Alho3, T. Asikainen6,
E. Asvestari1, M. Battarbee1, A. Binios3, A. Bosser3, T. Brito1, J. Envall2,
U. Ganse1, H. George1, J. Gieseler4, S. Good1, M. Grandin1, S. Haslam2,
H.-P. Hedman7, H. Hietala4, N. Jovanovic3, S. Kakakhel7, M. Kalliokoski1,
V. V. Kettunen3, T. Koskela1,4, E. Lumme1, M. Meskanen2, D. Morosan1,
M. Rizwan Mughal3, P. Niemel¨a3, S. Nyman3, P. Oleynik4, A. Osmane1,
E. Palmerio1, Y. Pfau-Kempf1, J. Peltonen4, J. Plosila7, J. Polkko2,
S. Poluianov6,8, J. Pomoell1, D. Price1, A. Punkkinen4, R. Punkkinen7,
B. Riwanto3, L. Salomaa7, A. Slavinskis3,9, T. S¨antti7, J. Tammi7,
H. Tenhunen7, P. Toivanen2, J. Tuominen7, L. Turc1, E. Valtonen4,
P. Virtanen4, T. Westerlund7
1University of Helsinki, Department of Physics, Helsinki, Finland
2Finnish Meteorological Institute, Space and Earth Observation Centre, Helsinki, Finland
3Aalto University, School of Electrical Engineering, Espoo, Finland
4University of Turku, Department of Physics and Astronomy, Turku, Finland
5University of Michigan, Department of Climate and Space Sciences and Engineering, Ann Arbor, USA
6University of Oulu, Space Climate Research Unit, Oulu, Finland
7University of Turku, Department of Future Technologies, Turku, Finland
8University of Oulu, Sodankyl¨a Geophysical Observatory (Oulu Unit), Oulu, Finland
9University of Tartu, Tartu Observatory, T˜oravere, Estonia
Key Points:
FORESAIL-1 mission measures energetic electron precipitation and solar ener-
getic neutral atom flux
We will demonstrate a cost-efficient de-orbiting and orbit manoeuvring technol-
ogy without propellants
The goal of the mission is to contribute significantly to the sustainable utilisation
of space
Corresponding author: Minna Palmroth,
arXiv:1905.09600v1 [] 23 May 2019
manuscript submitted to JGR: Space Physics
Today, the near-Earth space is facing a paradigm change as the number of new
spacecraft is literally sky-rocketing. Increasing numbers of small satellites threaten the
sustainable use of space, as without removal, space debris will eventually make certain
critical orbits unusable. A central factor affecting small spacecraft health and leading
to debris is the radiation environment, which is unpredictable due to an incomplete
understanding of the near-Earth radiation environment itself and its variability driven
by the solar wind and outer magnetosphere. This paper presents the FORESAIL-1
nanosatellite mission, having two scientific and one technological objectives. The first
scientific objective is to measure the energy and flux of energetic particle loss to the
atmosphere with a representative energy and pitch angle resolution over a wide range
of magnetic local times. To pave the way to novel model - in situ data comparisons,
we also show preliminary results on precipitating electron fluxes obtained with the
new global hybrid-Vlasov simulation Vlasiator. The second scientific objective of the
FORESAIL-1 mission is to measure energetic neutral atoms (ENAs) of solar origin.
The solar ENA flux has the potential to contribute importantly to the knowledge of so-
lar eruption energy budget estimations. The technological objective is to demonstrate
a satellite de-orbiting technology, and for the first time, make an orbit manoeuvre
with a propellantless nanosatellite. FORESAIL-1 will demonstrate the potential for
nanosatellites to make important scientific contributions as well as promote the sus-
tainable utilisation of space by using a cost-efficient de-orbiting technology.
1 Introduction
Unprecedented numbers of new spacecraft are now being launched into Earth or-
bit to satisfy the growing demand from the scientific, commercial, and military sectors.
Most of these new spacecraft need to survive in the radiation belts (RBs; van Allen
& Frank, 1959), which are regions of trapped energetic charged particles. The RBs
are critical in terms of space weather, as the radiation ages the spacecraft and deteri-
orates hardware. All new satellites contribute to the already existing large number of
orbital debris, if they are not actively removed at the end of the mission. This section
introduces the state of the art in the three scientific and technological objectives of
the FORESAIL-1 mission: measurements of energetic particle precipitation and solar
energetic neutral atoms (ENAs), and de-orbiting technologies.
1.1 State of the Art: Electron precipitation observations
The RBs are produced by a complex balance of particle source and loss processes
that vary both temporally and spatially (e.g., Tverskoy, 1969; Walt, 1996; Chen et
al., 2007; Shprits et al., 2008; Thorne, 2010). Significant variations in electron fluxes
occur over various time scales as a function of both energy and distance, driven by
solar-magnetospheric interactions and internal magnetospheric processes (e.g., Li et al.,
1997; Elkington et al., 2003; Reeves et al., 2003; Shprits et al., 2006; Baker & Kanekal,
2008). Effective losses from the outer radiation belts consist of 1) loss through the
outer edge of the magnetosphere (magnetopause shadowing (e.g., West et al., 1972;
Ukhorskiy et al., 2006; Saito et al., 2010; Matsumura et al., 2011; Turner et al., 2014)),
2) radial outward displacement of the electrons due to waves (Mann et al., 2016), and
weakening of the Earth’s magnetic field (the Dst effect (McIlwain (1966); Kim &
Chan (1997); Millan & Thorne (2007)), and 3) wave-particle interactions resulting in
scattering of particles into the loss cone (Kennel & Petschek, 1966; Thorne & Kennel,
1971; Thorne, 1974). There are no comprehensive estimates about which of these
processes is most important during different conditions, while it is clear that particle
losses play a central part in regulating the RBs.
manuscript submitted to JGR: Space Physics
The wave-particle interactions leading to losses from the RBs vary on time scales
ranging from 100 milliseconds to several minutes (Millan & Thorne, 2007). Balloon
experiments have historically been the earliest method to determine this loss category
by measuring the X-rays from bremsstrahlung radiation induced by the interaction of
precipitating particles with the neutrals in the upper atmosphere (Barcus et al., 1973;
Pytte et al., 1976). Latest such observations are provided by the BARREL mission
(Millan & the BARREL Team, 2011; Woodger et al., 2015). All balloon missions are
constrained to balloon-reachable altitudes and thus only allow indirect observation of
the precipitating particles.
Energetic particle precipitation has also been observed from the ground, as pre-
cipitating electrons with energies over several tens of keV cause enhanced ionisation
in the ionospheric D-region at an altitude of about 90 km. Relative ionospheric opac-
ity meters (riometers) (Hargreaves, 1969) are ground-based passive radars measuring
the so-called cosmic noise absorption (Shain, 1951), which corresponds to absorbed
radio wave power in the ionosphere resulting from enhanced D-region electron den-
sity. Recently, interferometric riometry has been developed to produce all-sky maps
(e.g. McKay-Bukowski et al., 2015). Ground-based observations of energetic electron
precipitation can also be achieved using incoherent scatter radars, which can accu-
rately measure D-region electron density down to about 70 km altitude (e.g. Miyoshi
et al., 2015). However, the intrinsically indirect ground-based observations do not
allow inferring precipitating particle fluxes and energies, even using newly developed
approaches such as spectral riometry (Kero et al., 2014). Hence having direct mea-
surements of precipitating fluxes from space-borne instruments is critical for radiation
belt loss studies.
One of the first satellite missions to study energetic electron precipitation was
the Solar, Anomalous, and Magnetospheric Particle Explorer (SAMPEX, 1992 - 2012)
used in a number of studies (Li et al., 2001; Tu et al., 2010; Nakamura et al., 2000).
DEMETER microsatellite observed electron fluxes at energies between 70 keV and
2.5 MeV with high energy resolution (256 channels) on a 700 km orbit (Sauvaud et
al., 2006). These observations have been used to infer energetic electron precipitation
(Graf et al., 2009), although DEMETER viewed primarily trapped particles. The
main data set of direct measurements of precipitating energetic particles comes from
the Medium Energy Proton and Electron Detector (MEPED) onboard NOAA/POES
satellites (D. S. Evans & Greer, 2000). MEPED consists of two telescopes, the 0
telescope designed to measure precipitating particle fluxes and the 90telescope for
trapped particle fluxes, measuring electrons in three energy channels (>30 keV, >100
keV, and >300 keV) and protons in five energy channels (>30 keV, >80 keV, >250
keV, >800 keV, and >2.5 MeV). However, the NOAA/POES particle data suffer from
two issues. First, the 0telescope only partially views the bounce loss cone and
does not offer any angular resolution inside its viewcone leading to poor pitch angle
resolution. This leads to an underestimation of the precipitating fluxes (Rodger et
al., 2013). Second, the electron channels are affected by proton contamination; partly
corrected by a new data set by Asikainen & Mursula (2013).
Particle precipitation is a key element in magnetosphere–ionosphere–thermosphere
coupling, and therefore a crucial objective for research in numerical models, especially
as there is an increasing demand for space weather forecasting capabilities. The first
attempts to model precipitating particle fluxes relied on statistical patterns inferred
from satellite observations. McDiarmid et al. (1975) produced a model for precipitat-
ing electron flux within 0.15–200 keV as a function of magnetic local time (MLT) and
invariant latitude based on about 1100 passes of the ISIS 2 spacecraft. Using data mea-
sured by the Low Energy Electron experiment onboard the Atmosphere Explorer C
and D satellites, Spiro et al. (1982) parametrised precipitating electron energy flux and
average energy as a function of MLT, geomagnetic latitude, and geomagnetic activity
manuscript submitted to JGR: Space Physics
measured by the Kp and AE indices. One of the reference models for auroral-energy
electron precipitation is the Hardy et al. (1985) model, empirically derived by com-
piling several years of observations from the Defense Meteorological Satellite Program
and Satellite (DMSP) Test Program spacecraft. The Hardy model is parametrised by
the Kp index, and it is still used to provide precipitation input in the 50 eV–20 keV
energy range to state-of-the-art ionospheric models (e.g., Marchaudon & Blelly, 2015).
At higher energies (30 keV–1 MeV), the recent van de Kamp et al. (2016) model pro-
vides energy-flux spectra of precipitating electrons as a function of Lparameter and
geomagnetic activity rendered with the Ap index. This empirical model was obtained
from a statistical analysis of 11 years of NOAA/POES energetic electron precipitation
observations and is averaged across all MLTs in its present version.
Modelling particle precipitation using first-principle models is not easy, given
that many processes operating at spatial and temporal scales spanning many orders of
magnitude are at play in the inner magnetosphere (energisation, and loss-cone scatter-
ing processes, among others). The emergence of global kinetic magnetospheric codes
may enable addressing this issue in the near future. Recently, a preliminary run was
performed using the Vlasiator code (von Alfthan et al., 2014; Palmroth et al., 2018), in
which electrons were added as a modelled species during a substorm-time, polar-plane
global magnetospheric run throughout the magnetospheric simulation box. Figure 1
shows an example of these preliminary results of 0.1 60 keV electron precipitation
estimation obtained from the analysis of the nightside velocity distribution functions
of electrons at a single time step of this simulation. The top panel shows the differen-
tial number flux of precipitating electrons as a function of Lshell (blue shading), as
well as the mean precipitating energy (black line) in the same units as typical space-
craft data. The bottom panel shows the integral energy flux as a function of L. The
integral energy flux was obtained by multiplying the differential number flux by the
corresponding energies, and integrating across energies. The mean precipitating en-
ergy was calculated by dividing the integral energy flux by the total number flux (i.e.,
the differential number flux integrated across energies). The Hardy et al. (1985) model
predicts a maximum integral energy flux of the order of 108 109 keV cm2s1sr1
in the midnight sector, reached at geomagnetic latitudes comprised between 62and
69(translating into Lvalues between 4.5 and 7.8), depending on geomagnetic activity
given by the Kp index. The preliminary results from Vlasiator in Fig. 1 are therefore
in reasonable agreement with those values.
With the expanding human activity in space, it is increasingly important to
measure precipitating particle fluxes in situ and predict the precipitation by modelling,
in order to understand the Earth’s radiation environment. While previous satellite
missions have provided a plethora of observations of the physical processes within the
precipitation environment, none of the missions were designed specifically with a clear
focus on precipitation. A number of new cubesat missions recently launched or being
built focus on precipitation. These include, for example, the CSSWE mission (Kohnert
et al., 2011), the ELFIN mission (Shprits et al., 2018), the Firebird mission Crew et al.
(2016), and the AMICal Sat mission (Barthelemy et al., 2018). FORESAIL-1 will be a
complementary mission, improving the spatial and temporal resolution of precipitation
that may be offered by these missions together.
1.2 State of the Art: Solar energetic neutral atom observations
The energy budgets of the solar corona and solar eruptions are major unsolved
questions in solar physics. Coronal heating leads to an abundance of suprathermal
particles in the corona. Suprathermal ions are important for estimating the energy
budget of an eruption (e.g., Emslie et al., 2004), however, they do not produce mea-
surable amounts of electromagnetic radiation, and thus their abundance is difficult to
measure. Direct in-situ measurements of suprathermal ions will be provided by the
manuscript submitted to JGR: Space Physics
Figure 1. Preliminary results of Vlasiator modelling of electron precipitation. (top) Differen-
tial flux of precipitating electrons as a function of Lshell, in the same units as usually measured
by telescopes onboard spacecraft. The black line indicates the mean precipitating energy. (bot-
tom) Total precipitating energy flux as a function of Lshell.
manuscript submitted to JGR: Space Physics
recently launched NASA Parker Solar Probe mission, but only within the outermost
reaches of the solar corona. During solar eruptions these suprathermal ions can be
driven to participate in charge exchange processes with neutral atoms, resulting in the
formation of solar energetic neutral atoms (ENAs). So far, ENAs have been measured
only during one single event (Mewaldt et al., 2009) with the IMPACT/LET instrument
onboard the twin the Solar TErrestrial RElations Observatory (STEREO) spacecraft
close to the beginning of the mission. Even these results may be questionable, as they
have been disputed by Simnett (2011), who suggested the ENA observations could be
explained by an earlier precursor event, detected as an electron burst.
1.3 State of the Art: Space debris removal technologies
The number of spacecraft in low-Earth orbit (LEO) is rapidly increasing, as
technological and regulatory changes of launchers have allowed smaller satellites. These
so-called nanosatellites typically do not have propulsion systems requiring bulky or
volatile propellants for orbit control or de-orbiting, making them a significant source of
future orbital debris. Furthermore, if they cannot handle high-radiation environments,
these nanosatellites will fail early, thus on one hand contributing to existing debris
and on the other hand defeating potential plans for active deorbiting at the end of
mission. The sustainable use of the near-Earth space has become of great interest
(e.g., Bastida Virgili et al., 2016). To avoid low-Earth orbits becoming unusable in the
future, also nanosatellites should be removed at end-of-life, otherwise the amount of
space debris will increase exponentially due to collisions with bigger objects (Kessler
& Cour-Palais, n.d.; Klinkrad, 1993; Bradley & Wein, 2009; Bonnal et al., 2013),
creating a potential danger to all later space missions. International standards have
been developed to impose requirements on space missions to mitigate space debris
production (e.g., European Space Agency, 2014; for Standardization, 2011). Thus, it
is crucial to develop robust instruments for both controlling the small satellite orbits
as well as for removing them from orbit after the end of the mission.
Apart from technological solutions for reduction of space debris that are inherent
to the satellite design, efforts for third-party orbit removal techniques are ongoing. Ap-
proaches include spacecraft that perform automatic rendez-vous, attachment and joint
de-orbiting of larger space debris items. Using high-powered lasers (either ground- or
satellite-based) to exert radiation pressure, or directly ablate the surface material of
space debris (a so-called ”Laser Broom”) has been a research project in both civilian
(Bekey, 1997; Phipps et al., 2012) and military (Campbell, 2000) projects. Compact
propulsion methods possibly allowing de-orbiting of nanosatellites include miniaturised
pulsed plasma and Hall-effect thrusters, which have reached commercial technologi-
cal readiness, but still require propellants and a sizeable energy budget (Tummala &
Dutta, 2017). Photonic solar sails have been investigated for propellantless propulsion
and used successfully for interplanetary missions (Tsuda et al., 2013), as well as multi-
ple nanosatellite missions (Lappas et al., 2011) with mixed success. Meanwhile, electric
sails, in which electrically charged structures interact with the ion environment, have
been proposed (Janhunen, 2004), and suitable packages have been implemented for
nanosatellites, but successful experimental verification is still outstanding.
2 Science goals
2.1 Mission statement
FORESAIL-1 is the first nanosatellite mission designed to measure the energy-
dependent pitch angle spectra of the precipitating radiation belt particles, and solar
ENA flux. Further, FORESAIL-1 will demonstrate the effectiveness of the plasma
brake as a means of manipulating the spacecraft orbit in operation and lowering the
manuscript submitted to JGR: Space Physics
spacecraft altitude to speed up de-orbiting at the end of the mission, thus addressing
the sustainability of LEO space operations.
2.2 Science objectives
The FORESAIL-1 mission answers the following science questions: What are the
pitch-angle and energy signatures of precipitation events as a function of MLT? How
is the precipitation pitch-angle and energy distribution dependent on geomagnetic
activity and driving from the solar wind? Thus, the FORESAIL-1 mission aims to
perform precise directional measurements of electron and proton precipitation, as well
as the energy spectrum and particle fluxes over a wide energy range (tens of keV to
approximately one MeV). The spacecraft orbit shall drift in magnetic local time to
allow determining precipitation as a function of MLT. The time resolution needs to be
good enough to describe the lower bound on precipitation budget due to wave-particle
interactions between chorus waves and electrons. Combining several measurements of
at least three energy channels into a full pitch-angle resolution throughout the LEO
region, with a time resolution of at least 15 s, will enable research of loss processes
from the RBs.
To understand the energy budget of solar eruptions, the second science goal of
this mission is to measure solar ENAs. This requires novel observations in an energy
range well exceeding the magnetospheric ENA range. For this purpose, we use the
geomagnetic field as a filter of solar particles and measure the ENA flux, thus inferring
the flux that originates from the solar direction. Solar ENAs can only be measured
reliably at energies exceeding the ring-current ion energies.
2.3 Technological objectives
In addition to the science objectives outlined above, FORESAIL-1 has a techno-
logical objective to ensure the sustainable use of space and to set a precedent for main-
taining clean orbits. The objective is to test the plasma brake technology and achieve
at least a 100 km lowering of the spacecraft altitude at mid-to-high altitude LEO. The
consequences of this lowering of the orbit are 1) the orbital drift of the mission allowing
monitoring the science objectives as a function of MLT, and 2) demonstrating that
the technology can be used to de-orbit spacecraft. We will observe efficiency and per-
formance of the plasma brake during the experiment to determine general information
about the orbit lowering process. The success of the plasma brake experiment (and
thus the completion of the mission’s sustainability goals) is dependent on the reliable
operation of the avionics, making reliability a primary design driver for FORESAIL-1.
Radiation effects are identified as a major potential source of failures, hence radiation
hardening techniques are used in the design.
3 Requirements
The study of precipitating electrons from the RBs is intrinsically coupled to
the characteristic energy ranges of the electron seed populations there. Scientific and
operational requirements are as follows:
1. NOAA/POES, for which the energy resolution is (E2-E1)/E1= 3 (where E1and
E2are the upper and lower limits for consecutive integral channels), provides
the lower reference bar in terms of energy resolution (Evans & Greer, 2006).
The nominal energy resolution for FORESAIL-1 is 0.4 between the upper and
lower limits of a channel.
2. The particle detector must have a sufficient discriminating ability between elec-
trons and protons, such that the electron channel does not suffer from proton
manuscript submitted to JGR: Space Physics
contamination. For lower energy channels there is no contamination, while in
the higher energy channels we allow a small background, however, this should
be so small that the electrons are discernible.
3. The orbit must drift to cover several MLTs.
4. For electron precipitation and solar ENA measurements, the orbital altitude
should lie between 400 km and 800 km.
5. The electron pitch angle resolution should include at least three bins measured
every 15 s.
6. The mission profile must allow for the use of the plasma brake for orbital and
altitude control.
7. The mission profile must provide the ability for daily updates of measurements
of orbital parameters to assess the effect of the plasma brake, once activated.
In order to achieve the above scientific requirements, the spacecraft spin axis must
be oriented with 3accuracy, with spin of 4.00 ±0.04 rpm. Attitude determination
system must supply the magnetic field vector with 1accuracy and the satellite position
must be known with 5-km accuracy. There must be a 1 kbs1data downlink.
4 Description of the Mission
4.1 General Concept
FORESAIL-1 is a nanosatellite mission in LEO designed to answer the science
objectives outlined in Section 2. The payload consists of a PArticle TElescope (PATE)
and a Plasma Brake (PB). PATE will measure energetic electrons in the energy range
80 800 keV as well as H+ions (protons) and neutral atoms in the energy range
0.310 MeV. PB consists of a tether that will be used to lower the spacecraft altitude.
The spacecraft is constrained under the CubeSat 3U standard to fit the two payloads.
4.2 Mission Timeline and Orbit
The manufacturing of the FORESAIL-1 payload and spacecraft started in 2018
and the manufacturing and integration will continue throughout 2019 until launch.
The spacecraft is scheduled to be launched in 2020 into a Sun-synchronous orbit at
or lower than 700 km altitude. Available launch opportunities are sought in 2019.
Following the successful launch of the mission, the 1-month commissioning phase is
scheduled to start immediately. After the commissioning phase, the mission’s primary
science phase is scheduled for 4 months at the initial Sun-synchronous orbit. After
the primary phase, the plasma brake will be applied to lower the spacecraft by more
than 100 km such that 1) the plasma braking force is fully demonstrated and 2) from
the lowered orbit the spacecraft will naturally de-orbit after 25 years. The lowering
of the orbit will inject the spacecraft into a drifting polar orbit, allowing precipitation
measurements in different MLTs. Following the successful orbital manoeuvring of the
spacecraft, the science phase continues with detecting particle precipitation in the
drifting orbit for at least 12 years. After this will be the ENA measurement phase.
There is a possibility of an extended science phase that will be scheduled depending
on the health of the spacecraft.
4.3 Spacecraft Conjunctions
FORESAIL-1 can be used in conjunction with various other spacecraft in or-
der to determine the origin of precipitating particles observed. Spacecraft that can
provide context to the FORESAIL-1 observations include the Solar and Heliospheric
Observatory (SOHO), STEREO, the Advanced Composition Explorer (ACE), Wind,
DSCOVR, the Geostationary Operational Environmental Satellites (GOES), and the
manuscript submitted to JGR: Space Physics
Parker Solar Probe. These missions directly monitor solar wind conditions, coronal
mass ejections and solar energetic particles.
The conditions encountered by FORESAIL-1 will depend strongly on the state of
the other regions in the magnetosphere. Simultaneous observations from satellites such
as Cluster, the Magnetospheric Multi-Scale mission, the Time History of Events and
Macroscale Interactions during Substorms spacecraft (THEMIS), and the Geomag-
netic Tail Lab (GEOTAIL) will be invaluable for understanding the global context in
which the FORESAIL-1 measurements are made. In particular, when located in the
relevant region, these spacecraft can monitor substorms occurring in the magnetotail
and the associated fast earthward plasma flows which are also the sources of the en-
ergetic particle precipitations to be measured by FORESAIL-1, in addition to the RB
source. They can also provide information about the wave activity in the magneto-
sphere, which will be key to interpreting the FORESAIL-1 observations. In the regions
closer to Earth, data from the recently-launched Arase mission in the radiation belts
will be of particular importance. Direct complementary observations to FORESAIL-
1 will be provided by the POES satellites, which will provide precipitating particle
data at similar energy ranges, however these data are often problematic and require
4.4 Coordinated Ground-Based Observations
Coordinated observation will also be possible with ground-based instrumenta-
tion. Whenever suitable conjunctions with riometer chains such as the Finnish chain
operated by Sodankyl¨a Geophysical Observatory or the Canadian chain (NORSTAR)
take place, it will be possible to compare energetic particle precipitation patterns to
those inferred from cosmic noise absorption measurements. The special case of the
Kilpisj¨arvi Atmospheric Imaging Receiver Array (KAIRA) is of particular interest, as
this instrument, which may be used as a multibeam, multifrequency riometer could
allow to finely study energetic precipitation along the FORESAIL-1 overpass. Indeed,
KAIRA can provide 1 s time resolution observations of cosmic noise absorption with
beams of 1030width, depending on the considered frequency, which is accurate
enough to study, e.g., individual patches of pulsating aurora with KAIRA (Grandin et
al., 2017), suggesting that it may be possible to study mesoscale (<100 km) structures
in energetic precipitation using FORESAIL-1 and KAIRA conjunctions.
During conjunctions, it can also prove valuable to combine FORESAIL-1 data
with observations of other space weather manifestations. For instance, auroral precip-
itation can be inferred from observations by all-sky imagers such as the MIRACLE
network in Fennoscandia, and the geomagnetic context of FORESAIL-1 measurements
can be given more accurately during conjunctions with ground-based magnetometer
networks as well as by making use of polar cap convection maps derived from Super
Dual Auroral Radar Network observations (SUPERDARN). Finally, FORESAIL-1
precipitation data could prove useful in studies focusing on electron density enhance-
ments in the ionosphere using, e.g., ionosondes or incoherent scatter radars such as
the European Incoherent Scatter radars (EISCAT). The latter ones also enable the
study of effects such as ionospheric Joule heating or ion outflow, as they measure
ion and electron temperatures and ion line-of-sight velocity, and when in a specific
configuration they also allow estimating electric fields (Nygr´en et al., 2011).
5 Payloads
The FORESAIL-1 mission will carry two science payloads, the Particle Telescope
(PATE) and the Plasma Brake (PB).
manuscript submitted to JGR: Space Physics
Figure 2. a) Mechanical design concept of PATE. b) Schematic of the anti-coincidence (AC)
and main detector (D) stack of each telescope.
5.1 Particle Telescope (PATE)
PATE measures energetic electrons, H+ions (protons) and neutral H atoms.
The targeted energy range of hydrogen measurement is 0.310MeV, which covers the
energies above the typical ring-current proton energies. This is to avoid the neutral
hydrogen background originating from the interaction of the ring current with the
geocorona. The primary energy range for electrons is 80 800 keV. In addition, the
instrument is sensitive to electrons at higher and lower energies in channels, where
reliable particle identification cannot be performed, but especially the high-energy
integral channel is still valuable since the contamination from heavier species comes
only from relativistic protons, which have low fluxes compared to relativistic electrons
in typical conditions.
PATE consists of two particle telescopes with identical stacks of detectors (see
Fig. 2). The longer Telescope 1 is directed along the long axis of the spacecraft,
that is, perpendicular to the rotation axis, and thus, it scans the directions in a plane
perpendicular to the rotation axis of the spacecraft. The shorter Telescope 2 is directed
along the rotation axis, so it can maintain a stable orientation. When the rotation axis
is pointed towards the Sun, the telescope is able to measure the neutral hydrogen flux
from the Sun. Note that the instrument itself does not determine the hydrogen charge
state but relies on the geomagnetic field as a rigidity filter and on the measurement
of angular distribution to disentangle the neutral flux from the solar direction. The
motivation for the use of longer collimator in Telescope 1 is to improve the pitch angle
angular resolution to better than 10 degrees for the scanning telescope.
Both telescopes have a mechanical collimator defining the aperture, consisting
of an aluminum tube housing 18 (12) (500 µm Al + 500 µm Ta) collimator rings in
Telescope 1 (Telescope 2), followed by a stack of silicon detectors, D1, D2, and D3,
measuring the energy losses of the particles in adjacent layers. The measured signals
allow the determination of particle energy and the identification of particle species
(electron / H). The thicknesses of the D detectors are 20 µm, 350 µm, and 350 µm,
respectively. The D1 detectors have a bias voltage of 5 V, while the other ones are
biased at 70 V. D1 and D2 detectors are segmented so that the central elements have
diameters of 5.2 mm while the total active-area diameters of both are 16.4 mm. D3
has a single active area of 16.4 mm diameter. Both detector stacks are covered at the
bottom of the collimator by two thin (nominally 0.5µm each, 1 mm apart) Ni foils
preventing low-energy (<250 keV) ions and soft (<500eV) photons from entering the
detector stack. Each aperture is limited from above by an annular anti-coincidence
manuscript submitted to JGR: Space Physics
detector AC1 (300 µm thickness) with a circular hole of 14.0 mm diameter in the middle
and and an outer active-area diameter of 33.8 mm. Another single-element circular
anti-coincidence detector AC2 (350µm thickness, 33.8 mm active-area diameter) at
the bottom of the stack flags particles penetrating the whole D detector stack. Note
that while the AC2 is operated in anti-coincidence with the D detectors for the nominal
energy range of the instrument (particles stopping in the D stack), PATE still analyses
the D detector pulse heights for particles triggering AC2 but not AC1 to provide
integral flux channels above the nominal energy ranges. The distances from upper
surface to upper surface in the detector stack AC1–D1–D2–D3–AC2 are 2.5 mm, 2.0
mm, 2.5 mm, and 2.5 mm. The lower Ni foil is 3.2 mm above the upper surface of
Signal processing is based on continuous sampling and digitization of the analog
signals and on digital filtering and pulse height analysis. The signal processing board
contains 16 Analog to Digital Converters (ADCs) and a Microsemi ProAsic3L Field
Programmable Gate Array (FPGA) handling the signal processing for both telescopes.
The signal sampling rate in each ADC channel is 10 MHz, and the digital streams are
processed by the FPGA, which is running at 40 MHz. The logic analyzes the incoming
digital data streams, detects pulses, and identifies the particle for each valid coincidence
event, counting and rejecting from further analysis any events not matching validity
criteria. Valid particle events are then counted in separate counters based on their
detection time, species and measured energy, forming the bulk of the science data
of the instrument. The electron (hydrogen) spectrum consists of seven (ten) energy
channels, log-spaced in measured energy.
5.1.1 Instrument Performance
The performance of the PATE electron measurement has been simulated using
GEANT4 (Agostinelli et al., 2003). The simulation is performed for an isotropic
electron distribution launched from a (15-cm radius) spherical surface surrounding
PATE. The simulated pulse heights of all D-detector signals are analysed to separate
electrons and hydrogen (ions/ENAs) and to measure particle energies, as in the FPGA-
based on-board analysis. Particles producing a hit (i.e., an energy deposit >50 keV)
only in D2 are identified as electrons and particles producing a hit only in D1 are
identified as hydrogen. Hit levels can be set freely in the logic, but values lower than
50 keV in D1a, D1c, AC1 and AC2 (see Fig. 2) should not be used as the simulated
RMS noise levels in those pads are around 9–14 keV. While electrons are able to
produce hits in D1 as well, there is only a minor level of electron contamination in
the nominal hydrogen energy channels, which require the energy deposit in D1 to
exceed 110 keV. If more than two adjacent D layers detect a pulse, the Delta E – E
method is used for clean species separation. The geometric factor of the seven electron
channels as a function of electron energy for (the shorter) Telescope 2 is shown in Fig.
3 (Oleynik & Vainio, 2019). The high-energy tails of the response functions are due
to the inevitable scattering of electrons off the detectors and other structures inside
the telescope, which prevents the full energy of the electron to be absorbed in active
detector elements. The internal energy resolution of the instrument is much better for
hydrogen than for electrons and the response functions are close to boxcar type within
the nominal energy range.
The geometric factors of both telescopes are mostly determined by the uppper-
most collimator ring (with an inner diameter of 21.5 mm) and the hole of the AC1
detector (diameter of 14 mm), which are at a vertical separation of 12.0 cm (7.0 cm)
in Telescope 1 (Telescope 2). The nominal value of the geometric factor is 0.037 cm2sr
(0.11 cm2sr) for Telescope 1 (Telescope 2), but especially electrons have somewhat
lower values (see Fig. 3) due to scattering off the Ni foils and the D1 detector, causing
trajectories to miss the D2 detector. The nominal angular widths of the acceptance
manuscript submitted to JGR: Space Physics
Figure 3. The geometric factor of PATE as a function of energy for the seven electron energy
channels within the nominal energy range (80–800 keV), simulated using GEANT4 (Agostinelli
et al., 2003). The range above 800 keV is additionally covered by a penetrating particle flux
cones (half width at half the on-axis value) for the two telescopes are 4.6and 7.9,
respectively. The instrument can also be operated in a mode where only the central
segments of the D1 and D2 are included in the coincidence logic while the rim areas
are logically included in the anti-coincidence. This allows to decrease the geometric
factors of the telescopes by a factor of about 7. Any detector element can also be
switched off from the logic entirely.
5.1.2 Mass, power and telemetry budgets
The mass of PATE is 1.2kg, consisting of the instrument box and mechanical
support structures (435 g), detector and pre-amplified board housings (290 g), the
collimators (180 g), cables (115 g), and the rest (180 g), including the back-end elec-
tronics stack. The power budget for PATE is 2.5 W, half of which is consumed by the
FPGA, with an additional margin of 20%, mainly required to account for the final
FPGA power consumption.
The telemetry budget of PATE is summarized in Table 1. Spectral resolution of
the data products for both electrons and hydrogen is on average ∆E/E 40% within
the nominal energy ranges, which means that the spectral counter data consists of eight
(ten) differential and one (two) integral channels for electrons (hydrogen). The rotation
period, nominally 15 s, of the satellite equals the time resolution of the instrument.
This main time frame is further broken in 36 angular sectors for the rotating telescope
to provide the angular distribution measurement. Both telescopes deliver, in addition
to the spectral counter data, also pulse height data samples, which allow an accurate
in-flight calibration and health monitoring of the detection system.
manuscript submitted to JGR: Space Physics
Data source Data rate [bit/s] Data amount per day [kiB]
Rotating telescope 993 10500
Solar pointing telescope 60 633
Housekeeping 13 135
Total 1066 11268
Table 1. Summary of the PATE telemetry budget during science operations.
Figure 4. Operating principle of the PB: an electric current system and a net thrust exerted
on the negative tether by the plasma ram flow.
5.2 Plasma Brake (PB)
5.2.1 Operating Principle
The Plasma Brake instrument is designed to measure the Coulomb drag, i.e.,
the braking force caused by the ionospheric plasma ram flow to an electrically charged
tether (Figure 4). When interacting with the surrounding plasma, the negative tether
gathers positive ions, which tends to neutralize the tether. To maintain the charge, the
tether is connected by a high-voltage power system to a conducting surface (deployable
booms for FORESAIL-1) that acts as an electron sink to close the current system
through the plasma (Janhunen, 2011). The braking force can be measured in two
modes. One is to monitor the system spin rate while charging the tether synchronously
with the tether rotation (PB Measurement). The other is to maintain a constant
charging and monitor the spacecraft velocity and orbital elements (PB Mode).
We employ 1 kV voltage, which is the maximum without risking electron field
emission from micrometeoroid-struck parts of the tether wires. At this voltage, the
expected nominal plasma brake thrust per tether length is 58 nN/m when the tether is
perpendicular to the ram flow. This value is obtained by using Equation 1 of Janhunen
(2014) and assuming plasma density of 3·1010 m3, mean ion mass of 10 proton masses,
ram flow speed of 7.8 km/s and tether’s effective electric radius of 1 mm. The tether’s
collected ion current is small, nominally 30 µA for a 300 m long tether.
5.2.2 Design
The tether is deployed from a chamber (blue) by the centrifugal force affecting
the tether tip mass (gray button inside the red collar) (Figure 5). The mechanical
interface through the satellite side panel is provided by an anti-static collar (red) to
avoid triple junctions of plasma, insulator, and high voltage tether. The tip mass is
locked during the launch by two launch locks located on opposite sides of the tether
chamber opening. The tether reel (dark gray) and the stepper motor that turns it
(orange) are nested inside the tether chamber. The tether high-voltage contact is
manuscript submitted to JGR: Space Physics
Figure 5. PB payload structure.
through the conducting reel and the slider (copper brown) attached straight to the
high-voltage converter (orange) board behind the tether chamber.
The PB tether is made of thin metallic wires that are periodically bonded to
each to withstand the micrometeoroid and space debris collisions (e.g., Sepp¨anen et
al., 2011). The single wire thickness is a few tens of micrometers to minimise the ion
current to the tether and thus the mass of the power system and size of the electron
gathering surface.
The nominal FORESAIL-1 tether length is 300 meters requiring 20 Nms of the
total angular momentum to deploy it. The initial angular speed of 180 /s is sufficient
to provide a centrifugal force of 0.4 cN which would safely pull out a tip mass weighing
2.5 g without breaking the tether. The angular momentum is provided by several con-
secutive satellite spin-up and reel-out maneuvers to compensate for the decreasing spin
rate associated with the increasing moment of inertia. Since magnetorquers are used
for attitude control, the angular acceleration is low to avoid the tether winding around
the satellite. It also indicates that a considerable amount of time is needed to provide
the angular momentum. However, after measuring the Coulomb drag force with a few
tens of meters of the tether, the force can be used to spin up the satellite by modu-
lating the tether voltage in synchronization with the rotation (charging downstream
to increase the spin rate). After deploying the tether and taking PB measurements,
the PB mode will start by continuously charging the tether which in turn will lower
the orbit and start the satellite drift in MLT. When reaching 600-km altitude and
gaining a sufficient drift in MLT, the satellite will be prepared for PATE observations
– the tether will be discarded to allow the spin axis being pointed towards the Sun. It
can be done by attempting to reel in the tether which might break because it would be
partially broken by micrometeoroids. A broken tether would deorbit in a few months
thanks to its large area/mass ratio. If the tether does not break, the attitude control
system and/or PB itself will have to provide an angular momentum to compensate for
an increasing spin rate.
5.2.3 Mass, Power and Telemetry Budgets
The mass of the PB is 0.6 kg including a margin of 25%. The structure (frame,
tether chamber, and motor mounting shaft) contributes 67% to the mass budget.
The size of the system is 67×84×96 mm. For the two measurement modes of the
PB, the power budget for PB is 600 mW. For the tether reeling, the motor and the
controller consumes 7 W continuously. In case the reel-out power cannot be provided
continuously by the spacecraft, the operation can be done in stages. The telemetry
budget of the PB is summarised in Table 2. To reduce the overall telemetry budget,
manuscript submitted to JGR: Space Physics
Mode Data rate [bit/s] Data amount per day [kiB]
Reel-Out/In 128 1350
PB Mode 19 200
PB Measurement 256 2700
Standby 64 675
Table 2. Summary of the PB telemetry budget for the operation modes.
Subsystem Planned mass (g) Mass with contingency (g) Fraction (%)
OBC 80 96 2
EPS 926 1093 23
Magnetorquers 240 288 6
UHF Transceiver 80 96 2
Antennas 50 60 1
Structure 1033 1240 26
PATE 1000 1200 26
PB 600 660 14
Total 4009 4733 100
Table 3. FORESAIL-1 mass budget.
the long-duration routine PB mode uses a lower telemetry rate. Frequent housekeeping
data are not required because in the PB mode, altitude change over weeks to months
will show the success of experiment.
6 Spacecraft platform
The platform has been designed to accommodate the payloads and to provide
data, power and mechanical interfaces. The overall mission tree for the FORESAIL-1
is presented in Figure 6. Since one of the key technological drivers for the mission
is reliability, the avionics stack has several designs to this end. The avionics stack is
enclosed in a vault providing substantially better shielding than what is typically seen
on CubeSats (around 4 mm equivalent aluminium, instead of the more typical 2 mm),
thus enhancing system tolerance to total dose. Single-event effects will be mitigated
using dual cold redundancy, hardware overcurrent protection, minimization of the
software footprint, and systematic data integrity checks. Finally, the FORESAIL-
1 components will be submitted to radiation test campaigns; the radiation response
data will be made available publicly in order to benefit from the broader field of (small)
satellite technology and help other designers addressing this issue.
The avionics stack consists of the Electrical Power System (EPS) for power collec-
tion, storage and distribution, Communication System for telemetry, On-Board Com-
puter (OBC) for telemetry handling and mission and payload management and data
storage, and Attitude Determination and Control System (ADCS) for maintaining the
attitude modes during different operation phases. The mechanical structure satisfies
dimensional limitations of the CubeSat standard and ensures modular configuration
of the spacecraft’s subsystems. The configuration of the platform is shown in Figure
7. The total mass budget is shown in Table 3.
manuscript submitted to JGR: Space Physics
Figure 6. FORESAIL-1 mission product tree.
OBC EPS Battery
Figure 7. Spacecraft structure including the subsystems without shielding.
manuscript submitted to JGR: Space Physics
Parameter X,Y axes Z axis
Number of wire turns 184 952
Nominal current, mA 26.2 19.3
Dipole moment, Am2 0.2 0.2
Power, mW 51.7 56.5
Table 4. Quantitative parameters for the air core magneotorquers.
6.1 Platform subsystems
Each payload has different requirements for attitude. PATE needs to be oriented
towards the Sun, while the detector with longer collimator scans the environment
in the directions perpendicular to the Earth–Sun vector. The PB needs the spinning
control for deploying and maintaining the tension of the tether. While the PB does not
need specific pointing direction, its tether deployment introduces a mass distribution
change that will require a proper control of the spacecraft angular momentum.
The attitude determination and control system (ADCS) is divided into Atti-
tude Determination System (ADS) and Attitude Control System (ACS). The ADS is
equipped with gyroscope (L3GD20H), magnetometer (LIS3MDL), and in-house built
sun sensors (Noman et al., 2017). The angular velocity of the spacecraft is a crucial
parameter for the payload attitude modes, thus gyroscope is necessary. Sun sensors are
required for a precise sun pointing control, and magnetometers are needed for prop-
erly using with attitude control. The outputs will provide a full attitude information
through using an unscented Kalman filter algorithm on all sensors.
The ACS changes the orientation of the spacecraft by using magnetorquers.
There is a closed loop feedback, which ensures the maintenance of desired attitude
by repeating the torque command until the desired orientation is achieved. The mis-
sion requirements for PATE and PB impose specific constraints on the attitude control.
The spacecraft uses the following attitude control modes:
Detumble mode: To stabilize the spacecraft after deployment
Spin control mode: In order to deploy the PB tether, this mode spins up the
spacecraft preferably with the spin axis being aligned with the earth pole. After
deorbiting, spin down might be required during tether reel in if the tether does
not break.
Sun pointing mode: This mode continuously points Telescope 2 (at Y axis)
towards the sun while spinning in order for Telescope 1 (at Z-axis) to scan the
Magnetorquers are sufficient to provide all necessary control modes for the manipula-
tion of the attitude and angular velocity. They are designed in-house, in form of air
coils, so that they can be integrated to the solar panels (for the X and Y axes), and
on a small factor magnetorquer driver board for the Z axis. All axes have two mag-
netorquers connected in parallel. The designed magnetorquers are driven through a
custom-built coil driver to optimize either consumed power or the generated magnetic
moment. The quantitative parameters of air core magnetorquers are given in Table 4.
The EPS of the satellite consist of solar panels, power conditioning and power
distribution (Ali et al., 2014), (Mughal et al., 2014).The solar panels are mounted on
manuscript submitted to JGR: Space Physics
every long side of the satellite. The solar panels at other sides than where PATE is
located consist of 7 solar cells connected in series, whereas the panel at the PATE side
consists of 6 cells in order to provide space for the telescope. The power conditioning
consists of switching buck converters to convert the incoming solar power to the battery
voltage level. A perturb and observe based algorithm extracts the maximum power
from the solar cells. Each subsystem houses a linear regulator to effectively convert the
voltage down to a stable 3.3 V with a low ripple factor. The theoretical efficiency of
the EPS is above 85 %. All losses have been accounted for in the calculation of power
budget. The maximum input power is 7 W, whereas the average power consumption
in the nominal mode 3.7 W.
The on-board computer (OBC) is the satellite’s main computer responsible for
computational and data storage needs, running the ADCS algorithms, and storing
the telemetry and housekeeping data logs. A safety-critical ARM Cortex-R4 based
processor is selected as the central OBC of the spacecraft. The processor features 256
k data rapid-access memory and 3 MB of program flash. There are external flash
memory devices also interfaced with the processor. To facilitate possible faults due to
radiation, the OBC houses two cold redundant symmetric processors. Only one of the
processors is active and powered. The arbiter switches the control to the redundant
processor in case of failure. The OBC is responsible for operational work during the
PB and PATE operations and collects all relevant telemetry data for downlink.
The UHF transceiver onboard the FORESAIL-1 consist of CC1125 transceiver
with maximum output power of 15 dBm (30 mW); an external power amplifier RF5110G
(gain: 31.5 dB, maximum output power 34.5 dBm) to amplify the power to desired
1.5-watt power in the transmit chain. In the receive chain, it consist of a Low Noise
Amplifier and a bandpass filter.
7 Ground Segment and Operations
The primary ground station used for satellite operation is located in Aalto Uni-
versity Campus in Espoo, Finland. The ground station operates mainly as a radio
amateur satellite station and has capabilities for operation on radio amateur UHF,
VHF and S-bands. Due to its northern location the ground station has an average
link time to polar orbit up to 90 minutes per day for all passes above the horizon.
Ground station radio systems are built based on the Software Defined Radio (SDR)
architecture which facilitates easy satellite-specific customization. The ground station
infrastructure and equipment are designed and operated by students and serve also for
educational purposes for the Aalto University. The Aalto University ground station
also operates as the mission control centre.
Typically, each satellite pass provides 10-20 min of link time between satellite and
ground station. The data rate requirements for FORESAIL-1 in science mode require
the downlink data rate to be approximately 8 kbps. Since the data rate requirements
are not stringent in order to accomplish both the missions, the ground station at Aalto
University easily handles the data rate requirements.
8 Data Products
FORESAIL-1 data products are outlined in Table 5. The Aalto University ground
station is responsible for downlinking the L0/L1 data containing general spacecraft
housekeeping data, PATE raw data, PB housekeeping and mission log information.
These low level data products are shared forward using file-based web interface. For
PATE data processing, Level 1 raw data files combined to ADCS metadata and orbital
information are used to produce calibrated and geolocated measurements of the particle
manuscript submitted to JGR: Space Physics
Provider Data product Details
Mission List of data availability
Spacecraft Orbital data, Attitude Includes position and altitude required to
estimate the PB performance
PB Electric current in the tether
L0 telemetry stream
L1 raw data
L2 calibrated data
L3 derived products Plasma density
L2 and plasma density including position
L0 telemetry stream
L1 raw data
L2 calibrated data (fluxes)
L3 derived products L2 including location
Precipitation maps (varying time resolutions)
Angular distribution data (spin resolution)
Event catalogues (mission duration)
Table 5. FORESAIL-1 data products. PATE fluxes are given as function of energy, time and
pointing azimuth.
These data products will be used to demonstrate that the PB de-orbits the
satellite at a measurable amount, as a function of time. The PATE data will be used
to infer the precipitation energy spectrum in time and place, addressing the science
objectives. The science data are open for everyone, with the nominal rules-of-the-road
typical in the field of space physics.
9 Summary and Discussion
The increasing number of small satellites launched into Earth’s orbit raises timely
concerns about the sustainable use of space. Small, rapidly built and launched satellites
have a large future potential for scientific and commercial use. However, the satellites
will become debris sooner rather than later, if they have poor radiation tolerance and
if they are not de-orbited at the end of mission. The Finnish centre Of excellence
in REsearch of SustAInabLe space (FORESAIL) funded by the Academy of Finland
tackles this issue by focusing both on science of the near-Earth radiation environment
and on novel technological solutions related to building more resilient instruments and
debris removal.
The first nanosatellite designed and built by the Centre of Excellence, FORESAIL-
1, is a 3U cubesat operating at polar LEO orbit at and below 700 km which will produce
energy-dependent pitch angle spectra of electrons and protons that precipitate from
the RBs into the ionosphere. Further, it will measure energetic neutral atoms (ENAs)
from the Sun and test the PB technology to lower the spacecraft altitude and manage
its orbit in space.
Today, nanosatellites can address significant scientific questions. This requires
focused measurements and innovative technological approaches, as well as coordination
with the other spacecraft and facilities operating simultaneously. The FORESAIL-1
PATE instrument will make unprecedented and precise measurements of precipitating
electrons and protons with high temporal resolution. It will be able to discriminate
between electrons and protons, over a wide energy range (80 – 800 keV for electrons
manuscript submitted to JGR: Space Physics
and 0.3 – 10 MeV for protons and neutral atoms). The large coverage of the polar
cap at different orbital planes is achieved by operating the PB at the beginning of the
mission, which sets the spacecraft orbital plane to drift in magnetic local time. This is
the first time such a manoeuvre is attempted with a nanosatellite, and if successful, it
will open new avenues for controlling the orbits of propellantless spacecraft, expanding
their operational and observational ranges.
FORESAIL-1 is targeted to make important advances in radiation belt physics.
With simultaneous observations in the solar wind, magnetosphere and from the ground,
FORESAIL-1 will allow quantifying the role of solar wind and outer magnetospheric
driving, and the role of different plasma waves in the inner magnetosphere in the precip-
itation process. The primary science phase allows connecting precipitation signatures
and mechanisms to geomagnetic activity levels and solar wind conditions. Together
with the novel Vlasiator model, a global hybrid-Vlasov simulation (Palmroth et al.,
2018), it will be possible to tie precipitation measurements to global processes in the
outer magnetosphere for the first time, as Vlasiator begins its full six-dimensional oper-
ation before the launch of FORESAIL-1. The six dimensions refer to three dimensions
in the ordinary space and three in the velocity space to describe the full particle dis-
tribution function, which is used to infer the energy spectrum and pitch angle in time.
At the time of writing, Vlasiator allows already 2D electron precipitation calculations
(see Fig. 1), which are in reasonable agreement with previous estimates (Hardy et
al., 1985). Once the model is fully 6D, in situ observations of electron precipitation
can be directly compared to kinetic processes anywhere in the magnetosphere, without
limitations as to whether a spacecraft traverses particular regions. This unprecedented
plan will likely open new avenues in space physics.
Successful observations of ENAs from the Sun will allow estimating for the first
time comprehensively estimating the suprathermal coronal ion population indirectly.
This is the key knowledge for improved understanding of the charged particle accel-
eration processes at the CME-driven shock waves close to the Sun and of the CME
energy budget. The solar ENA flux has been measured only once in a very fortuitous
event (Mewaldt et al., 2009). Observing the ENA flux from the Sun on a regular basis
as a function of time and solar activity is unprecedented.
The demonstration of the altitude manoeuvre of FORESAIL-1 will bring po-
tential for future applications for the PB as a standard tool for removing satellites
from their orbits. This is in particular a compelling solution considering the possi-
bly upcoming regulations for including debris mitigation techniques in newly launched
FORESAIL-1 is at the forefront of scientific nanosatellites. The advances we have
made will be particularly important in demonstrating the usefulness of nanosatellites
in making relevant physics and discovery measurements (ENA), whose spatio-temporal
resolution could be brought to a new level using fleets of nanosatellites. Technological
solutions of FORESAIL-1 have particularly far reaching impact for future debris re-
moval solutions and orbit control. All these aspects are expected to pave the way for
the sustainable use of space.
The Finnish Centre of Excellence in Research of Sustainable Space, building and
launching three FORESAIL missions is funded through the Academy of Finland with
grant numbers 312351, 312390, 312358, 312357, and 312356.
We acknowledge The European Research Council for Starting grant 200141-
QuESpace, with which Vlasiator was developed, and Consolidator grant 682068-PRESTISSIMO
awarded to further develop Vlasiator and use it for scientific investigations. We
manuscript submitted to JGR: Space Physics
gratefully also acknowledge the Academy of Finland (grant numbers 138599, 267144,
309937, and 309939). We acknowledge the CSC – IT Center for Science Grand Chal-
lenge grant for 2018, with which the Vlasiator simulation run was carried out. Vlasia-
tor (, (Palmroth, 2019)) is distributed under
the GPL-2 open source license at (Palmroth &
the Vlasiator team, 2019). Vlasiator uses a data structure developed in-house (San-
droos, 2019), which is compatible with the VisIt visualization software (Childs et
al., 2012) using a plugin available at the VLSV repository. The Analysator software
(, (Hannuksela & the Vlasiator team, 2019)
was used to produce the presented figures. The run described here takes several ter-
abytes of disk space and is kept in storage maintained within the CSC IT Center for
Science. Data presented in this paper can be accessed by following the data policy on
the Vlasiator web site.
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... The use of CubeSats is one plausible option to realize innovative technology demonstration of tether systems in space since they are developed and launched at a much lower cost than conventional satellites. The authors of this paper currently develop 3U CubeSat FORESAIL-1 to realize the world's first space demonstration of Coulomb drag propulsion [6], [7]. FORESAIL-1 will demonstrate the plasma brake for deorbiting in the Low Earth Orbit (LEO). ...
... FORESAIL-1 is the first in the FORESAIL mission series developed by the Finnish Centre of Excellence for Sustainable Space [7]. The FORESAIL-1 layout is shown in Fig. 2. The satellite is planned to be launched into LEO in 2021. ...
... Operating principle of Coulomb drag propulsion. Source: Adapted from[7]. ...
... It is evident from the executive summary of ISS, which encapsulates the achievements of innovative research on the orbiting laboratory that had created a positive impact on the quality of life on Earth and the future scope of the interdisciplinary researches globally for creating an impact on scientific advancement. [1][2][3][4][5][6][7][8][9][10][11][12][13][14][15][16] Herein we conduct theoretical studies for providing the proof of the concept of space debris recycling and energy conversion system in a microgravity environment with an intention to carry out real-time experiments in the space platform through the multinational collaboration with wide scope and benefits to humanity. The emerging application of outer space for the progression of the standard of living and quest for wisdom has led to the growth of space debris or space junk in orbits where the large number of satellites is operational. ...
... Various reports and industrial engineering data on the usefulness of the prevailing debris extenuation methods are inadequate for a realistic conclusion. [4][5][6][7][8][9][10][11][12][13][14][15][16] The present trend of global space activities indicates that in the ensuing decades' huge volumes of space debris will be in orbit and there is a need to update the debris alleviation normal exercises to stimulate competent and active practices to better abate the risks from space debris for the benign maneuvers of future space missions. [4][5][6][7]14 The fact is that even if all the upcoming space launches are called off, the space debris that already exists will be tendering threats for numerous decades to come before all of them re-enter the earth and burn off. ...
Full-text available
The space debris management and alleviation in the microgravity environment is a dynamic research theme of contemporary interest. Herein, we provide a theoretical proof of the concept of a lucrative energy conversion system that is capable of changing the space debris into useful powders in the International Space Station (ISS) for various bids. A specially designed broom is adapted to collect the space debris of various sizes. An optical sorting method is proposed for the debris segregation in the ISS by creating an artificial gravitational field. It could be done by using the frame‐dragging effect or gravitomagnetism. An induction furnace is facilitated for converting the segregated metal‐scrap into liquid metal. A fuel‐cell aided water atomization method is proposed for transforming the liquid debris into metal powder. The high‐energetic metal powders obtained from the space debris could be employed for producing propellants for useful aerospace applications, and the silicon powder obtained could be used for making soil for fostering the pharmaceutical‐flora in the space lab in the future aiming for the scarce‐drug discoveries for high‐endurance health care management. The proposed energy conversion system is a possible alternative for the space debris extenuation and its real applications in orbiting laboratories through the international collaboration for the benefits to humanity.
... The Aalto-1 satellite is currently active in orbit as of June 2020, and the tether deployment has not been demonstrated yet. The upcoming missions ESTCube-2 (Iakubivskyi et al., in press) and FORESAIL-1 (Palmroth et al., 2019) are planning to deploy few-hundred-meter-long tethers. ...
... The initial task of the RU is to deploy the tether and keep it stretched during the entire mission. In previous missions, such as ESTCube-1 (Slavinskis et al., 2015;Lätt et al., 2014;Envall et al., 2014) and Aalto-1 (Khurshid et al., 2014), and upcoming missions, such as ESTCube-2 (Iakubivskyi et al., in press) and FORESAIL-1 (Palmroth et al., 2019), there was, and will be, an aluminium mass weighing a few grams at the end of the tether serving as a passive RU. In the scope of the MAT mission, the independent operation unit is required in order to control the spin plane and, therefore, to keep the intended trajectory; it also manages the spin rate in order to avoid the orbital Coriolis effect, which influences the spin rate dramatically while orbiting the Sun (Toivanen and Janhunen, 2012). ...
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We present a detailed mechanical and thermal analysis of a stand-alone nanospacecraft that performs asteroid flybys in the main asteroid belt (2.75 AU) and one Earth flyby at the end of the mission to return the gathered data. A fleet of such nanospacecraft (<10 kg) has been proposed as part of the Multi-Asteroid Touring mission concept, a nearly propellantless mission where the electric solar wind sail (E-sail) is used for primary propulsion. The fleet makes flybys of thus far poorly characterised asteroid populations in the main belt and downlinks scientific data during the returning Earth flyby. The spacecraft size is close to a three-unit cubesat with a mass of less than 6 kg. The spacecraft is designed for a 3.2-year round trip. A 20-km-long E-sail tether is used. A remote unit is attached to the tether’s tip and stowed inside the spacecraft before the E-sail commissioning. The remote unit is slightly smaller than a one-unit cubesat with a mass of approximately 750 g. With an electrospray thruster, it provides angular momentum during tether deployment and spin-rate management while operating the E-sail. The selection of materials and configurations is optimised for thermal environment as well as to minimise the mass budget. This paper analyses the main spacecraft and remote-unit architectures along with deployment and operation strategies from a structural point of view, and thermal analysis for both bodies.
... Literature review reveals that through fundamental research and development a few products and services derived from space station activities are entering the souk and furthering healthy and peaceful life on Earth [1], which is motivating us to propose more challenging research in ISS. It is evident from the executive summary of ISS, which encapsulates the achievements of innovative research on the orbiting laboratory that had created a positive impact on the quality of life on Earth and the future scope of the interdisciplinary researches globally for creating an impact on scientific advancement [1][2][3][4][5][6][7][8][9][10][11][12][13][14][15][16]. Herein we conduct theoretical studies for providing the proof of the concept of space debris recycling and energy conversion system in a microgravity environment with an intention to carry out real-time experiments in the space platform through the multi-national collaboration with wide scope and benefits to humanity. ...
... The growing utilization of outer space for the advancement of the standard of living and hunt for wisdom has led to the accretion of space debris or space junk in high-density satellites orbits. Various reports and industrial engineering data on the usefulness of the prevailing debris extenuation methods are inadequate for a realistic conclusion [4][5][6][7][8][9][10][11][12][13][14][15][16]. J.-C. ...
... The Coulomb drag generated by the tether should decrease the spacecraft altitude from 700 km to 500 km in half a year. A further planned mission that could serve as technological demonstrator of the plasma brake technology is FORESAIL-1, the first satellite designed by the Finnish centre Of excellence in REsearch of SustAInabLe space (FORESAIL) [165]. The 3U-CubeSat FORESAIL-1 should be launched in a polar orbit at an altitude of 700 km. ...
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The Electric Solar Wind Sail (E-sail) is an innovative propellantless propulsion system conceived by Pekka Janhunen in 2004 for use in interplanetary space. An E-sail consists of a network of electrically charged tethers maintained at a high voltage level by an electron emitter. The electrostatic field surrounding the E-sail extracts momentum from the incoming solar wind ions, thus giving rise to the generation of a continuous thrust. In a geocentric context, the same physical principle is also exploited by the plasma brake, a promising option for reducing the decay time of satellites in low Earth orbits after the end of their operational life. This paper discusses the scientific advances of both E-sail and plasma brake concepts from their first design to the current state of the art. A general description of the E-sail architecture is first presented with particular emphasis on the proposed tether deployment mechanisms and thermo-structural analyses that have been carried out over the recent years. The working principle of an E-sail is then illustrated and the evolution of the thrust and torque vector models is retraced to emphasize the subsequent refinements that these models have encountered. The dynamic behavior of an E-sail is also analyzed by illustrating the mathematical tools that have been proposed and developed for both orbital dynamics and attitude control. A particular effort is devoted to reviewing the numerous mission scenarios that have been studied to date. In fact, the extensive literature about E-sail-based mission scenarios demonstrates the versatility of such an innovative propulsion system in an interplanetary framework. Credit is given to the very recent studies on environmental uncertainties, which highlight the importance of using suitable control strategies for the compensation of solar wind fluctuations. Finally, the applications of the plasma brake are thoroughly reviewed.
... A number of new space start-ups were founded as an outcome of this project. The (former) Aalto satellites group members have started and joined a number of new missions, such as ICEYE SAR satellite constellation, Aalto-3, Reaktor Hello World, FORESAIL [69], and Comet Interceptor [70]. The Aalto-1 design has been beneficial in the space technology curriculum and a source of inspiration for new students in the space technology lab. ...
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The design, integration, testing and launch of the first Finnish satellite Aalto-1 is briefly presented in this paper. Aalto-1, a three-unit CubeSat, launched into Sun-synchronous polar orbit at an altitude of approximately 500 km, is operational since June 2017. It carries three experimental payloads: Aalto Spectral Imager(AaSI), Radiation Monitor (RADMON) and Electrostatic Plasma Brake (EPB). AaSI is a hyperspectral imager in visible and near-infrared (NIR) wavelength bands, RADMON is an energetic particle detector and EPB is a de-orbiting technology demonstration payload. The platform was designed to accommodate multiple payloads while ensuring sufficient data, power, radio, mechanical and electrical interfaces. The design strategy of platform and payload subsystems consists of in-house development and commercial subsystems. The CubeSat Assembly, Integration & Test (AIT) followed Flatsat-Engineering-Qualication Model (EQM)-Flight Model (FM) model philosophy for qualification and acceptance. The paper briefly describes the design approach of platform and payload subsystems, their integration and test campaigns and spacecraft launch. The paper also describes the ground segment & services that were developed by Aalto-1 team.
... Due to their acceptable prices, nanosatellites are used in many areas such as science experiments, environmental and climate monitoring, Earth observation, air and sea surveillance, global positioning system [3] [4] [5] [6] [7] . With rapid rise in applications of nanosatellites, many novel design techniques in the satellite subsystems, platform and payloads have also emerged [8] [9] [10] [11] [12] [13]. Nanosatellites provide economical access to space, hence a good choice for small research institutes and universities, due to small mass and size; as well as low power and relatively affordable cost. ...
CubeSats have become increasingly important in the last two decades and are playing a very important role in space industry, especially with Earth oriented nanosatellites. Earth oriented nanosatellites require more precise attitude control when there is a requirement for good pointing accuracy. In such missions passive control systems are not suitable due to their low accuracy that is why active control systems are used. Different types of actuators are available, but magnetic actuators are best suited for Low Earth Orbit (LEO) nanosatellites. When current moves through the wire, a magnetic field is generated which is used to generate torque. The torque generated makes it possible to control the attitude of the satellite in the desired direction. The best option for nanosatellites is to use the Printed or embedded magnetorquers due to its scalable, reconfigurable and modular approach. Printing the magnetorquer into the internal layers of printed circuit board (PCB) reduce the harness complexities and space constraints effectively. The optimized design of the printed Magnetorquer is selected by analyzing various parameters such as turn width, distance between two turns, applied voltage, external dimensions and internal dimensions. The design also take into consideration other parameters such as generated torque, torque to power ratio, magnetic field to current ratio, consumed power, rotation time, and thermal analysis by changing the optimizing variables and taking into account the most important key design drivers. The design selection and results of the analysis concerning the selection of optimized parameters are presented.
CubeSats are a class of miniaturized satellites that have become increasingly popular in academia and among hobbyists due to their short development time and low fabrication cost. Their compact size, lightweight characteristics, and ability to form a swarm enables them to communicate directly with one another to inspire new ideas on space exploration, space-based measurements, and implementation of the latest technology. CubeSat missions require specific antenna designs in order to achieve optimal performance and ensure mission success. Over the past two decades, a plethora of antenna designs have been proposed and implemented on CubeSat missions. Several challenges arise when designing CubeSat antennas such as gain, polarization, frequency selection, pointing accuracy, coverage, and deployment mechanisms. While these challenges are strongly related to the restrictions posed by the CubeSat standards, recently, researchers have turned their attention from the reliable and proven whip antenna to more sophisticated antenna designs such as antenna arrays to allow for higher gain and reconfigurable and steerable radiation patterns. This paper provides a comprehensive survey of the antennas used in 120 CubeSat missions from 2003 to 2022 as well as a collection of single-element antennas and antenna arrays that have been proposed in the literature. In addition, we propose a pictorial representation of how to select an antenna for different types of CubeSat missions. To this end, this paper aims is to serve both as an introductory guide on CubeSats antennas for CubeSat enthusiasts and a state of the art for CubeSat designers in this ever-growing field.
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Magnetorquer based attitude control system capable of attaining high spin rates and precise pointing control is required for a 3U CubeSat satellite FORESAIL-1. The satellite, developed by Finnish Centre of Excellence, needs to maintain a spin rate of 24 ∘/s and precise pointing of the spin axis toward the Sun for the particle telescope instrument, as well as to reach 130 ∘/s spin rate for the deployment of the plasma brake. Mission requirements analysis and attitude system requirements derivation are presented, followed by actuator trade-off and selection, with detailed design of the complete attitude control system, including air-cored type of magnetorquer actuators and their drivers, made of H-bridge and filtering components. The design is based on several theoretical and practical considerations with emphasis on high power efficiency, such as effects of parallel and serial magnetorquer connections, modelling the magnetorquers with equivalent circuit models for finding of suitable driving frequency and extrapolation methods for efficient dipole moment usage. In-house manufacturing process of magnetorquers, using a custom 3D-printer setup, is described. Lastly, the testing and verification are performed, by measuring the performance of the manufactured hardware, circuit simulations and attitude control simulations. It is shown that the manufactured attitude control system fulfils all system requirements. Simulations also confirm the capability to satisfy mission requirements.
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This paper reviews Vlasov-based numerical methods used to model plasma in space physics and astrophysics. Plasma consists of collectively behaving charged particles that form the major part of baryonic matter in the Universe. Many concepts ranging from our own planetary environment to the Solar system and beyond can be understood in terms of kinetic plasma physics, represented by the Vlasov equation. We introduce the physical basis for the Vlasov system, and then outline the associated numerical methods that are typically used. A particular application of the Vlasov system is Vlasiator, the world’s first global hybrid-Vlasov simulation for the Earth’s magnetic domain, the magnetosphere. We introduce the design strategies for Vlasiator and outline its numerical concepts ranging from solvers to coupling schemes. We review Vlasiator’s parallelisation methods and introduce the used high-performance computing (HPC) techniques. A short review of verification, validation and physical results is included. The purpose of the paper is to present the Vlasov system and introduce an example implementation, and to illustrate that even with massive computational challenges, an accurate description of physics can be rewarding in itself and significantly advance our understanding. Upcoming supercomputing resources are making similar efforts feasible in other fields as well, making our design options relevant for others facing similar challenges.
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CubeSats provide a cost effective means to perform scientific and technological studies in space. Due to their affordability, CubeSat technologies have been diversely studied and developed by educational institutions, companies and space organizations all over the world. The CubeSat technology that is surveyed in this paper is the propulsion system. A propulsion system is the primary mobility device of a spacecraft and helps with orbit modifications and attitude control. This paper provides an overview of micro-propulsion technologies that have been developed or are currently being developed for CubeSats. Some of the micro-propulsion technologies listed have also flown as secondary propulsion systems on larger spacecraft. Operating principles and key design considerations for each class of propulsion system are outlined. Finally, the performance factors of micro-propulsion systems have been summarized in terms of: first, a comparison of thrust and specific impulse for all propulsion systems; second, a comparison of power and specific impulse, as also thrust-to-power ratio and specific impulse for electric propulsion systems.
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The objective of the Electron Losses and Fields INvestigation on board the Lomonosov satellite (ELFIN-L) project is to determine the energy spectrum of precipitating energetic electrons and ions and, together with other polar-orbiting and equatorial missions, to better understand the mechanisms responsible for scattering these particles into the atmosphere. This mission will provide detailed measurements of the radiation environment at low altitudes. The 400–500 km sun-synchronous orbit of Lomonosov is ideal for observing electrons and ions precipitating into the atmosphere. This mission provides a unique opportunity to test the instruments. Similar suite of instruments will be flown in the future NSF- and NASA-supported spinning CubeSat ELFIN satellites which will augment current measurements by providing detailed information on pitch-angle distributions of precipitating and trapped particles.
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VisIt is a popular open source tool for visualizing and analyzing data. It owes its success to its foci of increasing data understanding, large data support, and providing a robust and usable product, as well as its underlying design that fits today's supercomputing landscape. In this short paper, we describe the VisIt project and its accomplishments.
This study investigates the contribution of energetic (E > 30 keV) particle precipitation during a pulsating aurora event over Kilpisjärvi (L = 6.2) on 26 February 2014. It is based on the comparison of auroral blue-line emission (427.8 nm) data from an all-sky camera and cosmic noise absorption (CNA) data obtained from a multibeam experiment of the Kilpisjärvi Atmospheric Imaging Receiver Array (KAIRA) riometer. The data sets are compared for three KAIRA beams close to magnetic zenith. Results show a clear correlation between the measured CNA and the auroral blue-line emission during the event, for each beam. In addition, individual pulsations are observed for the first time in the cosmic noise absorption data measured by KAIRA and are found to be close-to-identical to the optical pulsations. This suggests that the modulation of electron precipitation during pulsating aurora takes place in a consistent way over a broad range of energies.
The influence of solar variability on the polar atmosphere and climate due to energetic electron precipitation (EEP) has remained an open question largely due to lack of a long-term EEP forcing dataset that could be used in chemistry-climate models. Motivated by this we have developed a model for 30–1000keV radiation belt driven EEP. The model is based on precipitation data from low-Earth orbiting POES satellites in the period 2002-2012 and empirically described plasmasphere structure, which are both scaled to a geomagnetic index. This geomagnetic index is the only input of the model and can be either Dst or Ap. Because of this, the model can be used to calculate the energy-flux spectrum of precipitating electrons from 1957 (Dst) or 1932 (Ap) onwards, with a time resolution of 1 day. Results from the model compare well with EEP observations over the period of 2002–2012. Using the model avoids the challenges found in measured datasets concerning proton contamination. As demonstrated, the model results can be used to produce the first ever >80 year long atmospheric ionization rate dataset for radiation belt EEP. The impact of precipitation in this energy range is mainly seen at altitudes 70-110km. The ionization rate dataset, which is available for the scientific community, will enable simulations of EEP impacts on the atmosphere and climate with realistic EEP variability. Due to limitations in this first version of the model, the results most likely represent an underestimation of the total EEP effect.
Since the discovery of the Van Allen radiation belts over 50 years ago, an explanation for their complete dynamics has remained elusive. Especially challenging is understanding the recently discovered ultra-relativistic third electron radiation belt. Current theory asserts that loss in the heart of the outer belt, essential to the formation of the third belt, must be controlled by high-frequency plasma wave–particle scattering into the atmosphere, via whistler mode chorus, plasmaspheric hiss, or electromagnetic ion cyclotron waves. However, this has failed to accurately reproduce the third belt. Using a data-driven, time-dependent specification of ultra-low-frequency (ULF) waves we show for the first time how the third radiation belt is established as a simple, elegant consequence of storm-time extremely fast outward ULF wave transport. High-frequency wave–particle scattering loss into the atmosphere is not needed in this case. When rapid ULF wave transport coupled to a dynamic boundary is accurately specified, the sensitive dynamics controlling the enigmatic ultra-relativistic third radiation belt are naturally explained.
We present initial dual spacecraft observations that for the first time both constrain the spatial scale size and provide spectral properties at medium energies of electron microbursts. We explore individual microburst events that occurred on 2 February 2015 using simultaneous observations made by the twin CubeSats which comprise the National Science Foundation (NSF) Focused Investigations of Relativistic Electron Bursts: Intensity, Range, and Dynamics (FIREBIRD II). During these microburst events, the two identically instrumented FIREBIRD II CubeSats were separated by as little as 11km while traversing electron precipitation regions in low-Earth orbit. These coincident microburst events map to size scales >120km at the equator. Given the prevalence of coincident and noncoincident events we conclude that this is of the same order of magnitude as that of the spatial scale size of electron microburst, an unknown property that is critical for quantifying their overall role in radiation belt dynamics. Finally, we present measurements of electron microbursts showing that precipitation often occurs simultaneously across a broad energy range spanning 200keV to 1MeV, a new form of empirical evidence that provides additional insights into the physics of microburst generation mechanisms.